Low noise turbine for geared turbofan engine

ABSTRACT

A method of designing a gas turbine engine comprises the steps of including a fan section with a fan. A turbine section is included having a first turbine and a high pressure turbine. A gear reduction is included between the fan and the first turbine, the gear reduction being configured to receive an input from the first turbine and to turn the fan at a lower speed than the first turbine in operation. The first turbine is designed to include a number of turbine blades in each of a plurality of rows of the first turbine, the first turbine blades operating at least some of the time at a rotational speed, and the number of blades and the rotational speed being such that the following formula holds true for at least one of the blade rows of the first turbine: (number of blades×speed)/60≥5500.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. patent application Ser. No.15/245,383, filed Aug. 24, 2016, which is a continuation of U.S. patentapplication Ser. No. 15/007,784, filed Jan. 27, 2016, which is acontinuation of U.S. patent application Ser. No. 14/996,544 filed Jan.15, 2016, which is a continuation-in-part of U.S. patent applicationSer. No. 14/795,931, filed Jul. 10, 2015, which is acontinuation-in-part of U.S. patent application Ser. No. 14/248,386,filed Apr. 9, 2014, which is a continuation-in-part of InternationalApplication No. PCT/US2013/020724 filed Jan. 9, 2013 which claimspriority to U.S. Provisional Application No. 61/592,643, filed Jan. 31,2012. U.S. patent application Ser. No. 14/248,386 further claimspriority to U.S. Provisional Application No. 61/884,660 filed Sep. 30,2013.

BACKGROUND

This application relates to the design of a turbine which can beoperated to produce noise to which human hearing is less sensitive.

Gas turbine engines are known, and typically include a fan deliveringair into a compressor. The air is compressed in the compressor anddelivered downstream into a combustor section where it was mixed withfuel and ignited. Products of this combustion pass downstream overturbine rotors, driving the turbine rotors to rotate.

Typically, there is a high pressure turbine rotor, and a low pressureturbine rotor. Each of the turbine rotors includes a number of rows ofturbine blades which rotate with the rotor. Typically interspersedbetween the rows of turbine blades are vanes.

The low pressure turbine can be a significant noise source, as noise isproduced by fluid dynamic interaction between the blade rows and thevane rows. These interactions produce tones at a blade passage frequencyof each of the low pressure turbine stages, and their harmonics.

The noise can often be in a frequency range to which humans are verysensitive. To mitigate this problem, in the past, a vane-to-blade ratioof the fan drive turbine has been controlled to be above a certainnumber. As an example, a vane-to-blade ratio may be selected to be 1.5or greater, to prevent a fundamental blade passage tone from propagatingto the far field. This is known as acoustic “cut-off.”

However, acoustically cut-off designs may come at the expense ofincreased weight and reduced aerodynamic efficiency. Stated another way,if limited to a particular vane to blade ratio, the designer may berestricted from selecting such a ratio based upon other characteristicsof the intended engine.

Historically, the low pressure turbine has driven both a low pressurecompressor section and a fan section. More recently, a gear reductionhas been provided such that the fan and low pressure compressor can bedriven at different speeds.

SUMMARY

In a featured embodiment, a method of designing a gas turbine enginecomprises the steps of including a fan section with a fan, the fanincluding at least one fan blade. The fan section is designed to achievea low fan pressure ratio less than about 1.45, wherein the low fanpressure ratio is measured across a fan blade alone. A turbine sectionhas a first turbine and a second turbine. A gear reduction is includedbetween the fan and the first turbine, and includes an epicycle geartrain having a gear reduction ratio of greater than about 2.5:1. Thegear reduction is configured to receive an input from the first turbineand to turn the fan at a lower speed than the first turbine inoperation. The first turbine is designed to achieve a pressure ratiogreater than about 5:1. The first turbine includes an inlet having aninlet pressure, and an outlet that is prior to any exhaust nozzle andhaving an outlet pressure. The pressure ratio of the first turbine is aratio of the inlet pressure to the outlet pressure. The first turbine isdesigned to further include a number of turbine blades in each of aplurality of rows of the first turbine, the first turbine bladesoperating at least some of the time at a rotational speed, and thenumber of blades and the rotational speed being such that the followingformula holds true for at least one of the blade rows of the firstturbine: (number of blades×speed)/60≥5500. The rotational speed is anapproach speed in revolutions per minute, taken at an approachcertification point as defined in Part 36 of the Federal AirworthinessRegulations. The gas turbine engine is designed to produce 15,000 poundsof thrust or more.

In another embodiment according to the previous embodiment, the formularesults in a number greater than 6000.

In another embodiment according to any of the previous embodiments, theformula results in a number less than or equal to about 10000.

In another embodiment according to any of the previous embodiments, theformula results in a number less than 7000.

In another embodiment according to any of the previous embodiments, theformula holds true for a majority of the blade rows of the firstturbine.

In another embodiment according to any of the previous embodiments, theformula results in a number greater than 6000.

In another embodiment according to any of the previous embodiments, theformula results in a number less than or equal to about 10000.

In another embodiment according to any of the previous embodiments, theformula results in a number less than 7000.

In another embodiment according to any of the previous embodiments, amid-turbine frame is arranged between the second turbine and the firstturbine.

In another embodiment according to any of the previous embodiments, acompressor section is configured to drive air along core flowpath, and aplurality of bearing systems is configured to support the first turbineand the second turbine. The mid-turbine frame includes airfoilspositioned in the core flowpath and is configured to support at leastone of the bearing systems.

In another embodiment according to any of the previous embodiments, thesecond turbine has two stages.

In another embodiment according to any of the previous embodiments, afirst compressor is included, and a shaft is configured to be driven bythe first turbine. The gear reduction is arranged intermediate the firstcompressor and the shaft.

In another embodiment according to any of the previous embodiments, thesecond turbine has two stages.

In another embodiment according to any of the previous embodiments, theengine is designed to achieve a bypass ratio greater than ten (10). Thefan is designed to have a low corrected fan tip speed less than about1150 ft/second, wherein the low corrected fan tip speed is an actual fantip speed in ft/second at an ambient temperature divided by [(Tambient °R)/(518.7° R)]^(0.5).

In another embodiment according to any of the previous embodiments, thefan section is designed for cruise.

In another embodiment according to any of the previous embodiments, theformula holds true for all of the blade rows of the first turbine.

In another embodiment according to any of the previous embodiments, theformula results in a number greater than 6000.

In another embodiment according to any of the previous embodiments, theformula results in a number less than or equal to about 10000.

In another embodiment according to any of the previous embodiments, theformula results in a number less than 7000.

In another embodiment according to any of the previous embodiments, theformula does not hold true for all of the blade rows of the firstturbine.

In another featured embodiment, a method of designing a gas turbineengine comprises the steps of including a fan section with a fan, thefan having at least one fan blade. The fan section is designed toachieve a low fan pressure ratio less than about 1.45, wherein the lowfan pressure ratio is measured across a fan blade alone. A turbinesection has a first turbine and a second turbine. A gear reduction isbetween the fan and the first turbine and includes an epicycle geartrain having a gear reduction ratio of greater than about 2.5:1. Thegear reduction is configured to receive an input from the first turbineand to turn the fan at a lower speed than the first turbine inoperation. The first turbine is designed to achieve a pressure ratiogreater than about 5:1, the first turbine including an inlet having aninlet pressure, and an outlet that is prior to any exhaust nozzle andhaving an outlet pressure. The pressure ratio of the first turbine is aratio of the inlet pressure to the outlet pressure. The first turbine isdesigned to further include a number of turbine blades in each of aplurality of rows of the first turbine, and the turbine blades of thefirst turbine operating at least some of the time at a rotational speed,and the number of blades and the rotational speed being such that thefollowing formula holds true for at least one of the blade rows of thefirst turbine: (number of blades×speed)/60≥5500. The gas turbine engineis designed to produce 15,000 pounds of thrust or more.

In another embodiment according to the previous embodiment, the formularesults in a number less than 7000.

In another embodiment according to any of the previous embodiments, theformula holds true for a majority of the blade rows of the firstturbine.

In another embodiment according to any of the previous embodiments, theformula results in a number less than 7000.

In another embodiment according to any of the previous embodiments, amid-turbine frame is arranged between the second turbine and the firstturbine.

In another embodiment according to any of the previous embodiments, acompressor section is configured to drive air along core flowpath, and aplurality of bearing systems is configured to support the first turbineand the second turbine, wherein the mid-turbine frame includes airfoilspositioned in the core flowpath and is configured to support at leastone of the bearing systems.

In another embodiment according to any of the previous embodiments, afirst compressor is included, and a shaft is configured to be driven bythe first turbine. The gear reduction is arranged intermediate the firstcompressor and the shaft.

In another embodiment according to any of the previous embodiments, theengine is designed to achieve a bypass ratio greater than ten (10),wherein the fan section is designed for cruise, and wherein the fan isdesigned to achieve a low corrected fan tip speed less than about 1150ft/second, wherein the low corrected fan tip speed is an actual fan tipspeed in ft/second at an ambient temperature divided by [(Tambient °R)/(518.7° R)]^(0.5).

In another embodiment according to any of the previous embodiments, theformula holds true for all of the blade rows of the first turbine.

In another featured embodiment, a method of designing a turbine sectioncomprises the steps of including a low pressure turbine designed toachieve a pressure ratio greater than about 5:1. The low pressureturbine includes an inlet having an inlet pressure, and an outlet thatis prior to any exhaust nozzle and having an outlet pressure. Thepressure ratio of the low pressure turbine is a ratio of the inletpressure to the outlet pressure. The low pressure turbine is furtherdesigned to include a number of turbine blades in each of a plurality ofrows of the low pressure turbine, a majority of the turbine blades ofthe low pressure turbine operating at least some of the time at arotational speed, and the number of blades and the rotational speedbeing such that the following formula holds true for at least one of theblade rows of the low pressure turbine: (number ofblades×speed)/60≥5500. The rotational speed is an approach speed inrevolutions per minute, taken at an approach certification point asdefined in Part 36 of the Federal Airworthiness Regulations.

Although the different examples have the specific components shown inthe illustrations, embodiments of this disclosure are not limited tothose particular combinations. It is possible to use some of thecomponents or features from one of the examples in combination withfeatures or components from another one of the examples.

These and other features disclosed herein can be best understood fromthe following specification and drawings, the following of which is abrief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a gas turbine engine.

FIG. 2 shows another embodiment.

FIG. 3 shows yet another embodiment.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown), or an intermediate spool,among other systems or features. The fan section 22 drives air along abypass flowpath B while the compressor section 24 drives air along acore flowpath C for compression and communication into the combustorsection 26 then expansion through the turbine section 28. Althoughdepicted as a two-spool turbofan gas turbine engine in the disclosednon-limiting embodiment, it should be understood that the conceptsdescribed herein are not limited to use with two-spool turbofans as theteachings may be applied to other types of turbine engines includingthree-spool architectures.

The engine 20 generally includes a low speed spool 30 and a high speedspool 32 mounted for rotation about an engine central longitudinal axisA relative to an engine static structure 36 via several bearing systems38. It should be understood that various bearing systems 38 at variouslocations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through ageared architecture 48 to drive the fan 42 at a lower speed than the lowspeed spool 30. The high speed spool 32 includes an outer shaft 50 thatinterconnects a high pressure compressor 52 and high pressure turbine54. A combustor 56 is arranged between the high pressure compressor 52and the high pressure turbine 54. A mid-turbine frame 57 of the enginestatic structure 36 is arranged generally between the high pressureturbine 54 and the low pressure turbine 46. The mid-turbine frame 57further supports bearing systems 38 in the turbine section 28. The innershaft 40 and the outer shaft 50 are concentric and rotate via bearingsystems 38 about the engine central longitudinal axis A which iscollinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion.

The terms “low” and “high” as applied to speed or pressure for thespools, compressors and turbines are of course relative to each other.That is, the low speed spool operates at a lower speed than the highspeed spool, and the low pressure sections operate at lower pressurethan the high pressures sections.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than ten (10), the gearedarchitecture 48 is an epicyclic gear train, such as a star system, aplanetary gear system or other gear system, with a gear reduction ratioof greater than about 2.3:1 or greater than about 2.5:1. In onedisclosed embodiment, the engine 20 bypass ratio is greater than aboutten (10:1), the fan diameter is significantly larger than that of thelow pressure compressor 44, and the low pressure turbine 46 has apressure ratio that is greater than about 5:1. The low pressure turbine46 pressure ratio is a ratio of the pressure measured at inlet of lowpressure turbine 46 to the pressure at the outlet of the low pressureturbine 46 (prior to an exhaust nozzle). It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFCT’)”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.50 and, insome embodiments, is less than about 1.45. “Low corrected fan tip speed”is the actual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tambient deg R)/518.7){circumflex over( )}0.5]. The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second.

The use of the gear reduction between the low pressure turbine spool andthe fan allows an increase of speed to the low pressure compressor. Inthe past, the speed of the low pressure turbine has been somewhatlimited in that the fan speed cannot be unduly high. The maximum fanspeed is at its outer tip, and in larger engines, the fan diameter ismuch larger than it may be in lower power engines. However, a gearreduction may be used to free the designer from compromising lowpressure turbine speed in order not to have unduly high fan speeds.

It has been discovered that a careful design between the number ofrotating blades, and the rotational speed of the low pressure turbinecan be selected to result in noise frequencies that are less sensitiveto human hearing.

A formula has been developed as follows:

(blade count×rotational speed)/(60 seconds/minute)≥4000 Hz.

That is, the number of rotating blades in any low pressure turbinestage, multiplied by the rotational speed of the low pressure turbine(in revolutions per minute), divided by 60 seconds per minute (to putthe amount per second, or Hertz) should be greater than or equal to 4000Hz. In one embodiment, the amount is above 5500 Hz. And, in anotherembodiment, the amount is above about 6000 Hz.

The operational speed of the low pressure turbine as utilized in theformula should correspond to the engine operating conditions at eachnoise certification point currently defined in Part 36 or the FederalAirworthiness Regulations. More particularly, the rotational speed maybe taken as an approach certification point as currently defined in Part36 of the Federal Airworthiness Regulations. For purposes of thisapplication and its claims, the term “approach speed” equates to thiscertification point.

Although the above formula only needs to apply to one row of blades inthe low pressure turbine 26, in one embodiment, all of the rows in thelow pressure turbine meet the above formula. In another embodiment, themajority of the blade rows in the low pressure turbine meet the aboveformula.

This will result in operational noise to which human hearing will beless sensitive.

In embodiments, it may be that the formula can result in a range ofgreater than or equal to 4000 Hz, and moving higher. Thus, by carefullydesigning the number of blades and controlling the operational speed ofthe low pressure turbine (and a worker of ordinary skill in the artwould recognize how to control this speed) one can assure that the noisefrequencies produced by the low pressure turbine are of less concern tohumans.

This invention is most applicable to jet engines rated to produce 15,000pounds of thrust or more and with bypass ratios greater than about 8.0.

FIG. 2 shows an embodiment 200, wherein there is a fan drive turbine 208driving a shaft 206 to in turn drive a fan rotor 202. A gear reduction204 may be positioned between the fan drive turbine 208 and the fanrotor 202. This gear reduction 204 may be structured and operate likethe gear reduction disclosed above. A compressor rotor 210 is driven byan intermediate pressure turbine 212, and a second stage compressorrotor 214 is driven by a turbine rotor 216. A combustion section 218 ispositioned intermediate the compressor rotor 214 and the turbine section216.

FIG. 3 shows yet another embodiment 300 wherein a fan rotor 302 and afirst stage compressor 304 rotate at a common speed. The gear reduction306 (which may be structured as disclosed above) is intermediate thecompressor rotor 304 and a shaft 308 which is driven by a low pressureturbine section.

Although an embodiment of this invention has been disclosed, a worker ofordinary skill in this art would recognize that certain modificationswould come within the scope of this invention. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this invention.

What is claimed is:
 1. A method of designing a gas turbine enginecomprising the steps of: including a fan section with a fan, the fanincluding at least one fan blade; the fan section designed to achieve alow fan pressure ratio less than about 1.45, wherein the low fanpressure ratio is measured across a fan blade alone; further including aturbine section having a first turbine and a second turbine; furtherincluding a gear reduction between the fan and the first turbine, thegear reduction including an epicycle gear train having a gear reductionratio of greater than about 2.5:1, and the gear reduction beingconfigured to receive an input from the first turbine and to turn thefan at a lower speed than the first turbine in operation; the firstturbine designed to achieve a pressure ratio greater than about 5:1, thefirst turbine including an inlet having an inlet pressure, and an outletthat is prior to any exhaust nozzle and having an outlet pressure, andthe pressure ratio of the first turbine being a ratio of the inletpressure to the outlet pressure; and the first turbine designed tofurther include a number of turbine blades in each of a plurality ofrows of the first turbine, the first turbine blades operating at leastsome of the time at a rotational speed, and the number of blades and therotational speed being such that the following formula holds true for atleast one of the blade rows of the first turbine: (number ofblades×speed)/60≥5500; wherein the rotational speed is an approach speedin revolutions per minute, taken at an approach certification point asdefined in Part 36 of the Federal Airworthiness Regulations; and whereinthe gas turbine engine is designed to produce 15,000 pounds of thrust ormore.
 2. The method as recited in claim 1, wherein the formula resultsin a number greater than
 6000. 3. The method as recited in claim 2,wherein the formula results in a number less than or equal to about10000.
 4. The method as recited in claim 3, wherein the formula resultsin a number less than
 7000. 5. The method as recited in claim 1, whereinthe formula holds true for a majority of the blade rows of the firstturbine.
 6. The method as recited in claim 5, wherein the formularesults in a number greater than
 6000. 7. The method as recited in claim6, wherein the formula results in a number less than or equal to about10000.
 8. The method as recited in claim 7, wherein the formula resultsin a number less than
 7000. 9. The method as recited in claim 5, furtherincluding a mid-turbine frame arranged between the second turbine andthe first turbine.
 10. The method as recited in claim 9, furtherincluding a compressor section configured to drive air along coreflowpath, and a plurality of bearing systems configured to support thefirst turbine and the second turbine, wherein the mid-turbine frameincludes airfoils positioned in the core flowpath and is configured tosupport at least one of the bearing systems.
 11. The method as recitedin claim 10, wherein the second turbine has two stages.
 12. The methodas recited in claim 5, further including a first compressor, and a shaftconfigured to be driven by the first turbine, the gear reductionarranged intermediate the first compressor and the shaft.
 13. The methodas recited in claim 12, wherein the second turbine has two stages. 14.The method as recited in claim 5, further designing the engine toachieve a bypass ratio greater than ten (10), and wherein the fan isdesigned to have a low corrected fan tip speed less than about 1150ft/second, wherein the low corrected fan tip speed is an actual fan tipspeed in ft/second at an ambient temperature divided by [(Tambient °R)/(518.7° R)]^(0.5).
 15. The method as recited in claim 14, wherein thefan section is designed for cruise.
 16. The method as recited in claim1, wherein the formula holds true for all of the blade rows of the firstturbine.
 17. The method as recited in claim 16, wherein the formularesults in a number greater than
 6000. 18. The method as recited inclaim 17, wherein the formula results in a number less than or equal toabout
 10000. 19. The method as recited in claim 18, wherein the formularesults in a number less than
 7000. 20. The method as recited in claim1, wherein the formula does not hold true for all of the blade rows ofthe first turbine.